1. Field of the Invention
The present invention relates to a gas turbine engine for an aircraft and such and more particularly relates to a shroud segment of a turbine shroud for the gas turbine engine.
2. Description of the Related Art
A turbine of a gas turbine engine for an aircraft and such is provided with plural stages of turbine shrouds for suppressing influence of hot combustion gas on a turbine case. The turbine shrouds are directly exposed to the hot gas and thereby a severe thermal stress might be applied thereto. To avoid an excessive thermal stress, in common, each turbine shroud is segmented. Plural shroud segments are built up to form each turbine shroud.
Each shroud segment is provided with a back plate, as a main body thereof, formed in an arc shape. An outer surface of the back plate is supported by the turbine case. An inner surface of the back plate is integrally provided with a touching member for touching with rotating turbine blades, which is formed in a honeycomb shape or the like. The back plate is further provided with a first plate portion and a second plate portion.
Both side surfaces of the back plate are respectively provided with first sealing slots. The first sealing slots receive first spline seal plates for suppressing leakage of the hot combustion gas to the low-pressure turbine case.
Similarly, both side surfaces of the second plate portion are respectively provided with second sealing slots communicating with the first sealing slots. The second sealing slots also receive second spline seal plates for suppressing leakage of the hot combustion gas to the low-pressure turbine case.
Such a shroud segment with first and second spline seal plates effectively suppress the leakage of the hot combustion gas to the low-pressure turbine case so that excessive heating of the low-pressure turbine case is prevented.
A related art is discloses in Japanese Patent Application Laid-open No. H09-329003.